Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

EPPLER 377 AIRFOIL (e377-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 377 AIRFOIL (e377-il)
Reynolds number: 200,000
Max Cl/Cd: 94.31 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e377-il-200000-n5.txt
Download as CSV file: xf-e377-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 377 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.1542   0.09598   0.09245  -0.0272   0.8219   0.0126
  -7.250  -0.1452   0.09391   0.09023  -0.0277   0.7872   0.0130
  -7.000  -0.1361   0.09202   0.08826  -0.0286   0.7688   0.0134
  -6.750  -0.1238   0.09017   0.08634  -0.0304   0.7526   0.0137
  -6.500  -0.1102   0.08836   0.08445  -0.0325   0.7369   0.0139
  -6.250  -0.0954   0.08633   0.08234  -0.0348   0.7223   0.0140
  -5.750  -0.0607   0.08204   0.07794  -0.0407   0.7008   0.0141
  -5.500  -0.0405   0.07970   0.07555  -0.0442   0.6919   0.0142
  -5.250  -0.0196   0.07720   0.07299  -0.0475   0.6822   0.0142
  -5.000   0.0033   0.07462   0.07033  -0.0513   0.6735   0.0143
  -4.750   0.0268   0.07193   0.06757  -0.0548   0.6650   0.0143
  -4.500   0.0493   0.06853   0.06414  -0.0580   0.6575   0.0144
  -4.250   0.0594   0.06418   0.05977  -0.0572   0.6503   0.0149
  -4.000   0.0775   0.06157   0.05712  -0.0584   0.6430   0.0164
  -3.750   0.1041   0.05918   0.05466  -0.0619   0.6359   0.0192
  -3.500   0.1421   0.05704   0.05242  -0.0679   0.6290   0.0214
  -3.250   0.1867   0.05518   0.05042  -0.0754   0.6222   0.0219
  -3.000   0.2257   0.05277   0.04789  -0.0811   0.6157   0.0220
  -2.750   0.2605   0.05010   0.04509  -0.0853   0.6099   0.0221
  -2.500   0.2740   0.04584   0.04086  -0.0851   0.6039   0.0231
  -2.250   0.3009   0.04361   0.03853  -0.0871   0.5985   0.0248
  -2.000   0.3586   0.04285   0.03756  -0.0952   0.5920   0.0308
  -1.750   0.3979   0.04068   0.03523  -0.0992   0.5865   0.0310
  -1.500   0.4337   0.03837   0.03278  -0.1023   0.5813   0.0311
  -1.250   0.4645   0.03498   0.02931  -0.1051   0.5757   0.0317
  -1.000   0.4943   0.03276   0.02698  -0.1070   0.5709   0.0323
  -0.500   0.5623   0.02867   0.02264  -0.1114   0.5605   0.0261
  -0.250   0.6010   0.02701   0.02076  -0.1140   0.5559   0.0294
   0.000   0.6358   0.02522   0.01882  -0.1160   0.5508   0.0280
   0.250   0.6721   0.02338   0.01677  -0.1178   0.5459   0.0263
   0.500   0.7094   0.02130   0.01439  -0.1197   0.5418   0.0249
   0.750   0.7497   0.01825   0.01092  -0.1217   0.5371   0.0239
   1.000   0.7872   0.01479   0.00664  -0.1232   0.5328   0.0242
   1.250   0.8157   0.01414   0.00571  -0.1231   0.5285   0.0259
   1.500   0.8437   0.01371   0.00511  -0.1229   0.5238   0.0289
   1.750   0.8710   0.01358   0.00496  -0.1227   0.5187   0.0370
   2.000   0.8984   0.01339   0.00468  -0.1224   0.5144   0.0509
   2.250   0.9256   0.01343   0.00466  -0.1222   0.5099   0.0662
   2.500   0.9525   0.01357   0.00479  -0.1219   0.5047   0.0806
   2.750   0.9794   0.01364   0.00486  -0.1216   0.5004   0.0899
   3.000   1.0062   0.01361   0.00479  -0.1213   0.4959   0.0957
   3.250   1.0330   0.01363   0.00485  -0.1211   0.4907   0.1030
   3.500   1.0601   0.01368   0.00488  -0.1208   0.4863   0.1120
   3.750   1.0870   0.01377   0.00500  -0.1206   0.4816   0.1255
   4.000   1.1137   0.01384   0.00515  -0.1203   0.4763   0.1400
   4.250   1.1401   0.01394   0.00525  -0.1200   0.4717   0.1548
   4.500   1.1666   0.01406   0.00542  -0.1197   0.4662   0.1742
   4.750   1.1929   0.01417   0.00560  -0.1194   0.4602   0.1930
   5.000   1.2191   0.01430   0.00580  -0.1191   0.4544   0.2165
   5.250   1.2453   0.01442   0.00601  -0.1188   0.4472   0.2429
   5.500   1.2712   0.01455   0.00620  -0.1184   0.4407   0.2811
   5.750   1.2962   0.01432   0.00648  -0.1181   0.4330   0.6907
   6.250   1.3508   0.01441   0.00694  -0.1180   0.4173   1.0000
   6.500   1.3762   0.01463   0.00721  -0.1175   0.4084   1.0000
   6.750   1.4011   0.01488   0.00747  -0.1171   0.3992   1.0000
   7.000   1.4260   0.01512   0.00778  -0.1166   0.3883   1.0000
   7.250   1.4506   0.01539   0.00811  -0.1161   0.3764   1.0000
   7.500   1.4748   0.01569   0.00852  -0.1156   0.3630   1.0000
   7.750   1.4986   0.01603   0.00891  -0.1150   0.3477   1.0000
   8.000   1.5216   0.01645   0.00935  -0.1143   0.3297   1.0000
   8.250   1.5437   0.01695   0.00986  -0.1136   0.3052   1.0000
   8.500   1.5619   0.01787   0.01057  -0.1126   0.2597   1.0000
   8.750   1.5760   0.01931   0.01167  -0.1113   0.2050   1.0000
   9.000   1.5908   0.02064   0.01283  -0.1101   0.1679   1.0000
   9.250   1.6024   0.02226   0.01427  -0.1086   0.1284   1.0000
   9.500   1.6133   0.02382   0.01566  -0.1070   0.0983   1.0000
   9.750   1.6215   0.02549   0.01718  -0.1051   0.0708   1.0000
  10.000   1.6250   0.02742   0.01894  -0.1028   0.0449   1.0000
  10.250   1.6248   0.02933   0.02076  -0.1001   0.0275   1.0000
  10.500   1.6191   0.03146   0.02285  -0.0970   0.0153   1.0000
  10.750   1.6138   0.03390   0.02533  -0.0947   0.0094   1.0000
  11.000   1.6116   0.03639   0.02794  -0.0932   0.0076   1.0000
  11.250   1.6106   0.03901   0.03070  -0.0922   0.0068   1.0000
  11.500   1.6088   0.04192   0.03375  -0.0914   0.0062   1.0000
  11.750   1.6057   0.04512   0.03710  -0.0909   0.0058   1.0000
  12.000   1.6016   0.04856   0.04071  -0.0906   0.0055   1.0000
  12.250   1.5975   0.05213   0.04450  -0.0905   0.0053   1.0000
  12.500   1.5938   0.05572   0.04827  -0.0904   0.0052   1.0000
  12.750   1.5888   0.05957   0.05230  -0.0906   0.0051   1.0000
  13.000   1.5828   0.06364   0.05655  -0.0908   0.0050   1.0000
  13.250   1.5762   0.06791   0.06100  -0.0913   0.0049   1.0000
  13.500   1.5686   0.07238   0.06564  -0.0919   0.0049   1.0000
  13.750   1.5609   0.07706   0.07049  -0.0927   0.0048   1.0000
  14.000   1.5531   0.08192   0.07552  -0.0937   0.0047   1.0000
  14.250   1.5455   0.08690   0.08067  -0.0949   0.0047   1.0000
  14.500   1.5380   0.09200   0.08592  -0.0962   0.0046   1.0000
<< Back to EPPLER 377 AIRFOIL (e377-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 377 AIRFOIL (e377-il)